This invention relates to deployable radiators for spacecraft, especially satellites.
A typical communications satellite in earth orbit is shown in FIG. 1. Power is generated by solar panels 1, 2, and r.f. power is transmitted from the dishes 3 or 4. Typically for 1 kW r.f. power generated, something like 4 kW is generated as heat by electrical equipment. This heat must be rejected from the satellite and, to do this, it is conducted by heat pipes or heat transport loops such as capillary pumped loops to radiating surfaces of the satellite. The surfaces 5, 6 of the satellite which face North and South form the radiating surfaces, since these surfaces receive least heat from the sun. It will be remembered that in outer space, heat cannot be rejected by conduction or convection.
The increasing demands placed on communications satellites require more power to be generated with the consequent need to reject more heat. The area of the surfaces 5, 6 cannot be increased without severe packaging problems at launch, and a solution which has been turned to is the use of additional radiating surfaces which are stowed when the satellite is being launched and deployed when it is released (U.S. Pat. No. 5,117,901). The hinge between the additional radiating surface and the satellite is formed by short helical sections of a closed length of a pipe containing a working fluid which also contains a capillary evaporator, and forms a capillary pumped loop. However, the pipe makes several traverses across the panel, where it acts as a condenser, and a relatively heavy evaporator therefore has to be provided to provide the necessary pumping pressure, while the single closed path of the pipe makes it vulnerable to micro-meteorite strikes.
Another proposal for a deployable radiator is the use of a relatively long and wide flat tube forming a heat pipe, which is rolled up for launch and unrolled under the force of vapour pressure of a working fluid in the tube when the satellite is in orbit. The unrolled tube forms a radiator panel attached to a radiating surface of the satellite. However, in a heat pipe, unlike in a capillary pumped loop, the same surfaces are used for evaporation as are used for condensation, and this brings certain disadvantages.
The invention provides a deployable radiator for a spacecraft, which comprises a plurality of capillary evaporators, each of which is connected in a respective closed loop of pipe containing a working fluid, and the loops extending across a linking structure over at least part of their length in such a way that the linking structure together with the loops can be rolled up for launch, and unrolled for deployment.
The arrangement permits convenient stowage of the radiator, while being less prone to a micro-meteorite strike than if a single capillary pump loop was employed.
Advantageously the linking structure unrolls in use under the influence of the spring force of the rolled-up loops and, while this spring force may form the entire force for unrolling the linking structure, an extension mechanism may be provided for assisting with the unrolling. The linking structure may be formed from a layer of conducting material, which may be mounted on a layer of plastics material. Such a layer of conducting material would serve to spread the heat out. The linking structure may be a continuous panel, or could be mesh-like, or even slatted.
The capillary evaporators themselves may be secured to a face of the spacecraft.